With reference to FIG. 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 14 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
It is known for gas turbine engines to use air bled from a compressor to cool and seal hot parts of the engine, typically in the turbine. The air may be bled from either the outer or inner annulus line of the compressor, at the upstream or downstream face of any stage of rotor blades.
The air is bled from the compressor through off-take ports. Various different off-take port designs are known, such as ramped off-takes and diffusing slots. For example, FIG. 2 shows a typical off-take port 50 formed in the outer annular wall 52 of a compressor stage, located between a stator 54 and rotor 56 of the compressor stage. In this example, the off-take port 50 has a ramped form, intended to reduce the pressure loss across the off-take and thus improve the efficiency of the cooling air system. Another example of a known system is shown in EP-A-1136679. A still further example of a known system is shown in GB-A-1310401.
GB-A-2420155 discloses a gas turbine engine with a multi-stage compressor. A number of equi-angularly spaced off-take ports are formed through the inner annular wall of the compressor, at an axial location between the penultimate and final stages of the compressor. Therefore the off-take ports rotate with the rotor blades of the compressor. The off-take ports lead into off-take passages that are formed perpendicular to the gas path air flow in the compressor. The off-take feeds a downstream cooling system.